1. Field of the Invention
The present invention relates generally to a turbine blade, and more specifically to cooling of the tip region of the turbine blade.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, a hot gas flow is developed in the combustor from the burning of a fuel with compressed air from the compressor and then passed through a multiple staged turbine to produce mechanical power. In an aero engine, the mechanical power drives the rotor shaft that is connected to a bypass fan. In an industrial gas turbine engine, the rotor shaft is connected to an electric generator that will produce electrical power. In both engines, the engine efficiency can be increased by passing a higher temperature gas into the turbine. However, the turbine inlet temperature is limited to the material properties of the first stage turbine airfoils, these airfoils being the stator vanes and the rotor blades.
Complex internal airfoil cooling passages have been proposed to provide high levels of airfoil cooling using a minimal amount of cooling air. Higher turbine inlet temperatures are obtainable by providing improved airfoil cooling. Also, since the compressed air used to cool, these airfoils is taken from the compressor, the use of a minimal amount of compressor bleed off air for the airfoil cooling will also increase the engine efficiency.
Airfoil cooling is also important in increasing the life of the airfoils. Hot spots can occur on sections of the airfoils that are not adequately cooled. These hot spots can cause oxidation that will lead to shortened life for the airfoil. Blade tips are especially subject to hot spots since it is nearly impossible to total eliminate the gap between the rotating blade tip and the stationary shroud that forms the gap. Without any gas, blade tip rubbing will occur which leads to other problems. Because of the presence of the tip gap, the hot gas can flow through the gap and expose the blade tip surface to the extreme high temperatures of the gas flow. Therefore, adequate blade tip cooling is also required to reduce hot gas flow leakage and to control metal temperature in order to increase part life.
Airfoils surfaces exposed to the high temperature gas flow are typically coated with a thermal barrier coating or TBC in order to allow for even higher temperatures. As the TBC technology improves, more industrial gas turbine (IGT) blades are applied with a thicker or low conductivity TBC. Cooling flow demand has been gradually reduced. As a result, there is not sufficient cooling flow for the design to split the total cooling flow into two or three flow circuits and utilize the forward flowing serpentine cooling design. Serpentine flow cooling circuits provide higher cooling capabilities than several straight channels in the airfoil because the overall cooling passage length is increased due to the looping of the circuit up and down the airfoil. Cooling flow for the blade leading and trailing edges has to be combined with the mid-chord flow circuit to form a single 5-pass serpentine flow circuit. However, for the forward 5-pass serpentine flow circuit with total blade cooling flow back flow margin (BFM) may become a design issue.
FIG. 1 shows a prior art cross section view of a turbine blade with a 5-pass aft flowing serpentine cooling circuit for a first stage blade. in this cooling circuit, the tip section film cooling is achieved by bleed off cooling air, from the serpentine tip turns. Cooling air bleed off from the 5-pass serpentine flow circuit thus reduces the cooling performance for the serpentine flow circuit. FIG. 2 shows a prior, art first stage turbine blade that uses the cooling circuit of FIG. 1, and FIG. 3 shows a diagram view of the FIG. 1 serpentine flow cooling circuit.